CN112577070B - Low-resistance and high-efficiency scramjet engine thrust chamber integrated design method - Google Patents

Abstract

The invention discloses a low-resistance and high-efficiency scramjet thrust chamber integrated design method. The thrust chamber obtained by the design method comprises a combustion chamber wall surface, a cavity front edge wall surface, a cavity bottom wall surface, a first compression section wall surface, The wall surface of the second compression section, the wall surface of the first expansion section and the wall surface of the second expansion section; the profile line of the wall surface of the first compression section and the profile line of the wall surface of the second expansion section are all spline curves, and the profile line of the wall surface of the second compression section is the same as The profile of the wall of the first expansion section is an arc line with equal radii; the bottom wall of the cavity, the wall of the first compression section, between the walls of the second compression section, between the walls of the first expansion section, and between the walls of the second expansion section are connected smoothly. The engine combustion chamber and the nozzle in the thrust chamber are continuously transitioned, and the integrated design realizes the depth matching of the parameters of the combustion chamber and the nozzle inlet, effectively improving the thrust performance of the nozzle; effectively optimizing the thrust distribution of the thrust chamber components, eliminating shock waves, reducing Internal flow losses for improved engine performance.

Description

A low-resistance and high-efficiency scramjet thrust chamber integrated design method

technical field

The invention relates to the technical field of engines, in particular to a low-resistance and high-efficiency scramjet thrust chamber integrated design method.

Background technique

Hypersonic aircraft is a strategic high-tech technology that achieves high-speed penetration, 2-hour global arrival, and cheap access to space. Its development will change the future of war. It is the new commanding height of aerospace technology in the 21st century. The world's major countries are competing to develop related technologies and pose a new threat to our national security. As the best alternative power plant for hypersonic air-breathing flight, scramjet has become a research hotspot of various aerospace powers. As the core component of the scramjet, the performance of the supersonic combustion chamber directly determines the success or failure of the entire engine development. Due to the high flow velocity and short residence time, the scramjet usually adopts a concave cavity structure to stabilize the flame and organize the combustion. The high-temperature gas generated by the combustion in the combustion chamber is expanded and accelerated through the nozzle to form thrust.

In the current scramjet design, the combustion chamber and the nozzle are separately designed in sections, and there is a problem of mismatching flow parameters, which is not conducive to the improvement of the combustion efficiency and thrust efficiency of the engine; the cavity placed in the supersonic airflow is the inner part of the engine. The high-pressure combustion area in the cavity directly forms a force opposite to the propulsion direction on the wall of the trailing edge, and the negative thrust formed on the wall of the trailing edge is the main source of the resistance of the combustion chamber.

SUMMARY OF THE INVENTION

In view of the above-mentioned deficiencies in the prior art, the present invention provides a low-resistance and high-efficiency scramjet thrust chamber integrated design method. Through the integrated design, the engine combustion chamber and the nozzle are continuously transitioned to realize the combustion chamber and the nozzle. The depth matching of the inlet parameters effectively improves the thrust performance of the nozzle.

In order to achieve the above purpose, the present invention provides a low-resistance and high-efficiency scramjet thrust chamber, which includes a combustion chamber wall surface, a cavity front edge wall surface, a cavity bottom wall surface, and a first compression section wall surface connected in sequence along the engine flow direction. , the wall surface of the second compression section, the wall surface of the first expansion section and the wall surface of the second expansion section;

The profile line of the wall surface of the first compression section and the profile line of the wall surface of the second expansion section are both spline curves, and the profile line of the wall surface of the second compression section and the profile line of the wall surface of the first expansion section are arcs with equal radii Wire;

between the bottom wall of the cavity and the wall of the first compression section, between the wall of the first compression section and the wall of the second compression section, between the wall of the second compression section and the wall of the first expansion section, and between the wall of the first expansion section and the first expansion section. The walls of the two expansion sections are connected smoothly.

As a further improvement of the above technical solution, the profile of the wall of the first compression section is a second-order continuous derivable spline curve, and the first-order derivative, the second-order derivative and the concave curve of one end of the profile of the wall of the first compression section The first-order derivative and second-order derivative of the corresponding end of the cavity bottom wall are equal, and the first-order derivative and second-order derivative of the other end of the profile line of the first compression section wall are equal to the first-order derivative and second-order derivative of the corresponding end of the second compression section wall. .

As a further improvement of the above technical solution, the profile line of the front edge wall of the cavity and the profile line of the bottom wall of the cavity are both straight lines, and the profile line of the bottom wall of the cavity is parallel to the profile line of the combustion chamber wall. The profile line of the front edge wall of the cavity is vertically connected with the profile line of the bottom wall of the cavity and the profile line of the combustion chamber wall.

In order to achieve the above object, the present invention also provides an integrated design method for the above-mentioned low-resistance and high-efficiency scramjet thrust chamber, comprising the following steps:

Step 1. Given the depth of the front edge wall of the cavity and the total axial length of the cavity, the total axial length of the cavity is the sum of the bottom wall of the cavity, the axial length of the wall of the first compression section and the wall of the second compression section. ;

Step 2, carry out numerical simulation based on the air parameters of the engine inflow direction, the inlet size of the combustion chamber, the depth of the wall of the leading edge of the cavity and the total axial length of the cavity, and obtain the length of the recirculation zone in the cavity based on the numerical simulation results, That is, the axial length of the bottom wall of the cavity is obtained, and the sum of the axial lengths of the wall of the first compression section and the wall of the second compression section is obtained based on the total axial length of the cavity and the axial length of the bottom wall of the cavity;

Step 3, obtaining the throat height of the thrust chamber based on the air parameters in the incoming flow direction of the engine, wherein the throat is the connection between the wall surface of the second compression section and the wall surface of the first expansion section;

Step 4, select the second-order continuous derivable spline curve as the profile line of the first compression section wall, select the arc line as the profile line of the second compression wall surface, and obtain the profile line radius of the second compression wall surface based on the throat height , combined with the deflection angle of the second compression wall to obtain the axial length of the wall of the first compression section, the wall of the second compression section, and the curve configuration of the profile of the wall of the first compression section;

Step 5: Obtain the total axial length of the thrust chamber based on the structural constraints of the engine design, and obtain the axis of the expansion segment based on the total axial length of the thrust chamber, the axial length of the wall of the first compression section, and the axial length of the wall of the second compression section. The total axial length of the thrust chamber is the sum of the axial lengths of the wall surface of the first compression section, the wall surface of the second compression section, the wall surface of the first expansion section and the wall surface of the second expansion section, and the total axial length of the expansion section is the first The axial length of the wall surface of the expansion section and the wall surface of the second expansion section is the sum of the axial lengths of the wall surface of the first expansion section. The axial lengths of the walls of the second compression section are equal;

Step 6, based on the flow parameters of the throat, the total axial length of the expansion section, the height of the throat and the size constraints of the outlet of the thrust chamber, the axial length and profile configuration of the wall surface of the second expansion section are obtained.

As a further improvement of the above technical solution, in step 3, the throat height of the thrust chamber is obtained based on the air parameter of the incoming flow direction of the engine, specifically:

为发动机流量、T₀为燃烧室总温度、P₀为燃烧室压力、M为推力室喉部设计马赫数、R为气体常数、γ为比热比。where _h8 is the throat height of the thrust chamber;

is the engine flow rate, T ₀ is the total temperature of the combustion chamber, P ₀ is the combustion chamber pressure, M is the design Mach number of the thrust chamber throat, R is the gas constant, and γ is the specific heat ratio.

As a further improvement of the above technical solution, in step 4, the axial length of the wall surface of the first compression section, the wall surface of the second compression section, and the curve configuration of the profile of the wall surface of the first compression section are obtained in combination with the deflection angle of the second compression section wall. ,Specifically:

It is known that the profile of the second compression wall is a circular arc, the radius of the circular arc is R ₅ , the deflection angle is θ, and the end position of the circular arc is the throat of the thrust chamber, because in step 3 The throat height has been obtained, that is, the coordinates of an endpoint of the profile line of the second compression wall are obtained;

Combined with the radius R ₅ and the deflection angle θ, all the coordinates of the profile of the second compression wall can be obtained. Since the profile of the second compression wall intersects with the profile of the first compression wall, the profile of the wall of the first compression segment can be obtained. The coordinate position of an endpoint on the line;

At the same time, in steps 1 and 2, the depth of the front wall of the cavity and the axial length of the bottom wall of the cavity have been obtained, so all the coordinates of the profile of the bottom wall of the cavity can be obtained. The profile line of the first compression wall intersects, so to the coordinate position of the other end point of the profile line of the first compression segment wall;

On the premise that the coordinates of both ends of the profile line of the first compression section wall are known, and that the profile line is a second-order continuously derivable spline curve, the coordinates of all points on the profile line of the first compression section wall can be determined; that is, The axial lengths l ₅ and l ₆ of the wall of the first compression section, the wall of the second compression section, and the curved configuration of the profile of the wall of the first compression section can be obtained.

In order to achieve the above object, the present invention also provides a scramjet with the above-mentioned low-resistance and high-efficiency scramjet thrust chamber.

A low-resistance and high-efficiency scramjet thrust chamber integrated design method provided by the present invention has the following beneficial effects:

1. The combustion chamber and the nozzle of the engine are continuously transitioned, and the integrated design realizes the deep matching of the parameters of the combustion chamber and the nozzle inlet, and effectively improves the thrust performance of the nozzle;

2. Effectively optimize the thrust distribution of thrust chamber components, eliminate shock waves, reduce internal flow losses, and improve engine performance.

Description of drawings

In order to explain the embodiments of the present invention or the technical solutions in the prior art more clearly, the following briefly introduces the accompanying drawings that need to be used in the description of the embodiments or the prior art. Obviously, the accompanying drawings in the following description are only These are some embodiments of the present invention, and for those of ordinary skill in the art, other drawings can also be obtained according to the structures shown in these drawings without creative efforts.

Fig. 1 is the structural representation of thrust chamber in the prior art;

2 is a schematic structural diagram of a low-resistance and high-efficiency scramjet thrust chamber in an embodiment of the present invention;

3 is a schematic flowchart of a low-resistance and high-efficiency scramjet thrust chamber integrated design method in an embodiment of the present invention;

FIG. 4 is a schematic view of the dimensions in the process of the integrated design method of the thrust chamber of the low-resistance and high-efficiency scramjet in the embodiment of the present invention.

Description of reference numerals: combustion chamber wall 1, cavity front edge wall 2, cavity bottom wall 3, first compression section wall 4, second compression section wall 5, first expansion section wall 6, second expansion section wall 7 .

The realization, functional characteristics and advantages of the present invention will be further described with reference to the accompanying drawings in conjunction with the embodiments.

Detailed ways

The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention.

It should be noted that all directional indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present invention are only used to explain the relationship between various components under a certain posture (as shown in the accompanying drawings). The relative positional relationship, the movement situation, etc., if the specific posture changes, the directional indication also changes accordingly.

In addition, descriptions such as "first", "second", etc. in the present invention are only for descriptive purposes, and should not be construed as indicating or implying their relative importance or implicitly indicating the number of indicated technical features. Thus, a feature delimited with "first", "second" may expressly or implicitly include at least one of that feature. In the description of the present invention, "plurality" means at least two, such as two, three, etc., unless otherwise expressly and specifically defined.

In the present invention, unless otherwise expressly specified and limited, the terms "connected", "fixed" and the like should be understood in a broad sense, for example, "fixed" may be a fixed connection, a detachable connection, or an integrated; It can be a mechanical connection, an electrical connection, a physical connection or a wireless communication connection; it can be a direct connection or an indirect connection through an intermediate medium, and it can be the internal connection of two elements or the interaction between the two elements. unless otherwise expressly qualified. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific situations.

In addition, the technical solutions between the various embodiments of the present invention can be combined with each other, but must be based on the realization by those of ordinary skill in the art. When the combination of technical solutions is contradictory or cannot be achieved, it should be considered that the combination of technical solutions does not exist and is not within the scope of protection claimed by the present invention.

As shown in FIG. 2 , a low-resistance and high-efficiency scramjet thrust chamber disclosed in this embodiment includes a combustion chamber wall surface 1 , a cavity front edge wall surface 2 , and a cavity bottom wall surface 3 connected in sequence along the engine flow direction. , the first compression section wall surface 4, the second compression section wall surface 5, the first expansion section wall surface 6 and the second expansion section wall surface 7; the profile line of the first compression section wall surface 4 and the profile line of the second expansion section wall surface 7 are both Spline curve, the profile line of the wall surface 5 of the second compression section and the profile line of the wall surface 6 of the first expansion section are circular arcs of equal radius; between the bottom wall surface 3 of the cavity and the wall surface 4 of the first compression section, the first compression section The wall surface 4 and the wall surface 5 of the second compression section, the wall surface 5 of the second compression section and the wall surface 6 of the first expansion section, and the wall surface 6 of the first expansion section and the wall surface 7 of the second expansion section are connected smoothly.

Further preferably, the profile line of the wall surface 4 of the first compression stage is a second-order continuous derivable spline curve, and the first-order derivative and the second-order derivative of one end of the profile line of the wall surface 4 of the first compression stage correspond to the bottom wall surface 3 of the cavity The first-order derivative and second-order derivative of the end are equal, and the first-order derivative and second-order derivative of the other end of the profile line of the wall surface 4 of the first compression stage are equal to the first-order derivative and second-order derivative of the corresponding end of the wall surface 5 of the second compression stage.

It should be noted that in this embodiment, the profile line of the cavity front edge wall 2 and the profile line of the cavity bottom wall surface 3 are both straight lines, and the profile line of the cavity bottom wall surface 3 is parallel to the profile line of the combustion chamber wall surface 1. The profile line of the front edge wall surface 2 of the cavity is vertically connected with the profile line of the bottom wall surface 3 of the cavity cavity and the profile line of the combustion chamber wall surface 1 .

Based on the structure of the above-mentioned low-resistance and high-efficiency scramjet thrust chamber, this embodiment also discloses a low-resistance and high-efficiency scramjet thrust chamber integrated design method. Referring to Figures 3-4, the design method specifically includes the following step:

Step 1, the depth of the front edge wall of the cavity and the total axial length L ₁ of the cavity are given; wherein, the total axial length of the cavity is the axial length l ₃ of the bottom wall of the cavity and the axial length of the wall of the first compression section. The length sum of l ₄ and the axial length l ₅ of the wall of the second compression section, that is, L ₁ =l ₃ +l ₄ +l ₅ ;

Step 2: Numerical simulation is performed based on the air parameters in the incoming flow direction of the engine, the inlet size of the combustion chamber, the depth of the wall of the leading edge of the cavity, and the total axial length of the cavity, and the length of the recirculation zone in the cavity is obtained based on the numerical simulation results, That is, the value of the axial length l ₃ of the bottom wall of the cavity can be obtained, and the axial length of the axial length of the wall surface of the first compression section and the wall surface of the second compression section and the value of l ₄ +l ₅ can also be obtained in combination with step 1. ;

In step 3, the throat height of the thrust chamber is obtained based on the air parameters in the incoming flow direction of the engine, wherein the throat is the connection between the wall surface of the second compression section and the wall surface of the first expansion section, specifically:

为发动机流量、T₀为燃烧室总温度、P₀为燃烧室压力、M为推力室喉部设计马赫数、R为气体常数、γ为比热比；where _h8 is the throat height of the thrust chamber;

is the engine flow, T ₀ is the total temperature of the combustion chamber, P ₀ is the combustion chamber pressure, M is the design Mach number of the thrust chamber throat, R is the gas constant, and γ is the specific heat ratio;

Step 4, select the second-order continuous derivable spline curve as the profile line of the first compression section wall, select the arc line as the profile line of the second compression wall surface, and obtain the profile line radius of the second compression wall surface based on the throat height R5, through the given deflection angle θ of the second compression wall;

It is known that the profile of the second compression wall is a circular arc, the radius of the circular arc is R5, the deflection angle is θ, and the end position of the circular arc is the throat of the thrust chamber. The throat height is obtained, that is, the coordinates of one end point of the profile line of the second compression wall can be obtained. Combined with the radius R5 and the deflection angle θ, all the coordinates of the profile line of the second compression wall can be obtained. The profile line of a compression wall intersects, and then the coordinate position of an endpoint on the profile line of the first compression section wall can be obtained; at the same time, in steps 1 and 2, the depth of the front wall of the cavity and the axis of the bottom wall of the cavity have been obtained. Therefore, all the coordinates of the profile line of the bottom wall of the cavity can be obtained. Since the profile line of the bottom wall of the cavity intersects the profile line of the first compression wall, the distance to the other end of the profile line of the wall of the first compression section Coordinate location. When the coordinates of the two ends of the profile line of the wall of the first compression section are known, and the profile line is a second-order continuous derivable spline curve, the coordinates of all points on the profile line of the wall of the first compression section can be determined; that is, it can be obtained The axial lengths l ₅ and l ₆ of the wall of the first compression section, the wall of the second compression section, and the curved configuration of the profile of the wall of the first compression section are obtained.

Step 5: Obtain the total axial length of the thrust chamber based on the structural constraints of the engine design, and obtain the axis of the expansion section based on the total axial length of the thrust chamber, the axial length of the wall of the first compression section, and the axial length of the wall of the second compression section. The total axial length L ₂ of the thrust chamber is the axial length l ₄ of the wall surface of the first compression section, the axial length l ₅ of the wall surface of the second compression section, the axial length l ₆ of the wall surface of the first expansion section and The length sum of the axial length l ₇ of the wall surface of the second expansion section, that is, L ₂ =l ₄ +l ₅ +l ₆ +l ₇ ; the total axial length of the expansion section is the axial length l ₆ of the wall surface of the first expansion section and the _The length sum of the axial length _l7 of the wall surface of the second expansion section, the profile line of the wall surface of the first expansion section is a circular arc and its radius R6 is equal to the profile radius of the second compression wall surface, and the axis of the wall surface of the first expansion section is a circular arc. The axial length is equal to the axial length of the wall of the second compression section, that is, R ₅ =R ₆ , l ₅ =l ₆ ;

Step 6, based on the flow parameters of the throat, the total axial length of the expansion section L ₃ =l ₆ +l ₇ , the throat height h ₈ and the size constraint h ₇ of the thrust chamber outlet to obtain the axial length and profile of the wall of the second expansion section structure. Among them, the throat flow parameters include throat temperature and throat pressure, specifically:

where T is the throat temperature and p is the throat pressure.

In this embodiment, the Rao method of the maximum thrust nozzle design is used to obtain the axial length and profile configuration of the wall surface of the second expansion section, and the matching design of the profile line of the expansion section is completed. For the specific process, reference may be made to the document "Maurice J. Zurow, Joe D. Hoffman, Gas Dynamics Volume II, 1977, John Wiley & Sons, p164-169", and thus will not be repeated in this embodiment.

Based on the above-mentioned low-resistance and high-efficiency scramjet thrust chamber integrated design method, the low-resistance and high-efficiency scramjet thrust chamber in this embodiment can be obtained, in which the engine combustion chamber and the nozzle are continuously transitioned and integrated. It can effectively improve the thrust performance of the nozzle, and can effectively optimize the thrust distribution of the thrust chamber components, eliminate shock waves, reduce internal flow losses, and improve engine performance.

The above descriptions are only the preferred embodiments of the present invention, and are not intended to limit the scope of the present invention. Under the inventive concept of the present invention, the equivalent structural transformations made by the contents of the description and drawings of the present invention, or the direct/indirect application Other related technical fields are included in the scope of patent protection of the present invention.

Claims (3)

1. a low-resistance and high-efficiency scramjet thrust chamber integrated design method, the low-resistance and high-efficiency scramjet thrust chamber comprises a combustion chamber wall surface, a cavity leading edge wall surface, the bottom wall surface of the cavity, the wall surface of the first compression section, the wall surface of the second compression section, the wall surface of the first expansion section and the wall surface of the second expansion section; The profile line of the wall surface of the first compression section and the profile line of the wall surface of the second expansion section are both spline curves, and the profile line of the wall surface of the second compression section and the profile line of the wall surface of the first expansion section are arcs with equal radii Wire; between the bottom wall of the cavity and the wall of the first compression section, between the wall of the first compression section and the wall of the second compression section, between the wall of the second compression section and the wall of the first expansion section, and between the wall of the first expansion section and the first expansion section. The walls of the two expansion sections are connected smoothly; The profile line of the wall surface of the first compression section is a second-order continuous derivable spline curve, and the first-order derivative and the second-order derivative of one end of the profile line of the first compression section wall surface are the first-order derivative of the corresponding end of the cavity bottom wall surface. , the second-order derivatives are equal, and the first-order derivative and second-order derivative of the other end of the profile line of the first compression section wall are equal to the first-order derivative and second-order derivative of the corresponding end of the second compression section wall surface; The profile line of the front edge wall of the cavity and the profile line of the bottom wall of the cavity are both straight lines, the profile line of the bottom wall of the cavity is parallel to the profile line of the combustion chamber wall, and the profile line of the front edge wall of the cavity It is vertically connected to the profile line of the bottom wall of the cavity and the profile line of the combustion chamber wall; It is characterised in that the design method comprises the following steps: Step 1. Given the depth of the front edge wall of the cavity and the total axial length of the cavity, the total axial length of the cavity is the sum of the bottom wall of the cavity, the axial length of the wall of the first compression section and the wall of the second compression section. ; Step 2, carry out numerical simulation based on the air parameters of the engine inflow direction, the inlet size of the combustion chamber, the depth of the wall of the leading edge of the cavity and the total axial length of the cavity, and obtain the length of the recirculation zone in the cavity based on the numerical simulation results, That is, the axial length of the bottom wall of the cavity is obtained, and the sum of the axial lengths of the wall of the first compression section and the wall of the second compression section is obtained based on the total axial length of the cavity and the axial length of the bottom wall of the cavity; Step 3, obtaining the throat height of the thrust chamber based on the air parameters in the incoming flow direction of the engine, wherein the throat is the connection between the wall surface of the second compression section and the wall surface of the first expansion section; Step 4, select the second-order continuous derivable spline curve as the profile line of the first compression section wall, select the arc line as the profile line of the second compression wall surface, and obtain the profile line radius of the second compression wall surface based on the throat height , combined with the deflection angle of the second compression wall to obtain the axial length of the wall of the first compression section, the wall of the second compression section, and the curve configuration of the profile of the wall of the first compression section; Step 5: Obtain the total axial length of the thrust chamber based on the structural constraints of the engine design, and obtain the axis of the expansion segment based on the total axial length of the thrust chamber, the axial length of the wall of the first compression section, and the axial length of the wall of the second compression section. The total axial length of the thrust chamber is the sum of the axial lengths of the wall surface of the first compression section, the wall surface of the second compression section, the wall surface of the first expansion section and the wall surface of the second expansion section, and the total axial length of the expansion section is the first The axial length of the wall surface of the expansion section and the wall surface of the second expansion section is the sum of the axial lengths of the wall surface of the first expansion section. The axial lengths of the walls of the second compression section are equal; Step 6, based on the flow parameters of the throat, the total axial length of the expansion section, the height of the throat and the size of the outlet of the thrust chamber are constrained to obtain the axial length and profile configuration of the wall of the second expansion section; In step 3, the throat height of the thrust chamber is obtained based on the air parameter of the incoming flow direction of the engine, specifically:

In the formula,

is the throat height of the thrust chamber;

is the engine flow, T ₀ is the total temperature of the combustion chamber, P ₀ is the combustion chamber pressure, M is the design Mach number of the thrust chamber throat, R is the gas constant,

is the specific heat ratio; In step 6, the throat flow parameters include throat temperature and throat pressure, specifically:

where T is the throat temperature, p is the throat pressure, T ₀ is the total temperature of the combustion chamber, and P ₀ is the pressure of the combustion chamber. 2. The low-resistance and high-efficiency scramjet thrust chamber integrated design method according to claim 1, wherein in step 4, the wall surface of the first compression section and the second compression section are obtained in combination with the deflection angle of the second compression wall surface. The axial length of the wall, and the curved configuration of the profile of the wall of the first compression section, specifically: It is known that the profile of the second compression wall is a circular arc, and the radius of the circular arc is R ₅ , the deflection angle is θ , and the end position of the circular arc is the throat of the thrust chamber, because in step 3 The throat height has been obtained, that is, the coordinates of an endpoint of the profile of the second compression wall are obtained; Combined with the radius R ₅ and the deflection angle θ , all the coordinates of the profile of the second compression wall can be obtained. Since the profile of the second compression wall intersects with the profile of the first compression wall, the profile of the wall of the first compression segment can be obtained. The coordinate position of an endpoint on the line; At the same time, in steps 1 and 2, the depth of the front wall of the cavity and the axial length of the bottom wall of the cavity have been obtained, so all the coordinates of the profile of the bottom wall of the cavity can be obtained. The profile line of the first compression wall intersects, so to the coordinate position of the other end point of the profile line of the first compression segment wall; On the premise that the coordinates of both ends of the profile line of the first compression section wall are known, and that the profile line is a second-order continuously derivable spline curve, the coordinates of all points on the profile line of the first compression section wall can be determined; that is, The axial lengths l ₅ and l ₆ of the wall of the first compression section, the wall of the second compression section, and the curved configuration of the profile of the wall of the first compression section can be obtained. 3. A scramjet, characterized in that it has a low-resistance and high-efficiency scramjet thrust chamber designed by the design method of claim 1 or 2.

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